This invention relates generally to articles which are used in a high-temperature environment. More particularly, it relates to methods for protecting the articles from damage in such an environment.
Various types of materials, such as metals and ceramics, are used for components which may be exposed to a high-temperature environment. Aircraft engine parts represent examples of these types of components. Peak gas temperatures present in the turbine engine of an aircraft are maintained as high as possible for operating efficiency. Turbine blades and other elements of the engine are usually made of metal alloys which can resist the high-temperature environment, e.g., superalloys, which have an operating temperature limit of about 1000xc2x0 C.-150xc2x0 C. Operation above these temperatures may cause the various turbine elements to fail and damage the engine.
A variety of approaches have therefore been used to raise the operating temperature of the componentsxe2x80x94especially, metal components. For example, one approach involves the use of protective coatings on the surfaces of the components. The coatings are usually ceramic-based, and are sometimes referred to as thermal barrier coatings or xe2x80x9cTBCxe2x80x9ds.
Another approach (which may be used in conjunction with the TBC""s) calls for the incorporation of internal cooling :channels in the metal component, through which cool air is forced during engine operation. As an example, a pattern of cooling holes may extend from a relatively cool surface of a combustion chamber to a xe2x80x9chotxe2x80x9d surface which is exposed to gas flow at combustion temperatures of at least about 10000xc2x0 C. (The drawings described below will serve to illustrate this concept.) The technique is sometimes referred to as xe2x80x9cdiscrete hole film coolingxe2x80x9d. Cooling air, usually bled off from the engine""s compressor, is typically bypassed around the engine""s combustion zone and fed through the cooling holes to the hot surface. The ratio of the cooling air mass flux (the product of air velocity times density) to the mass flux of the hot gas flowing along the hot surface (e.g., a combustion product) is sometimes referred to as the xe2x80x9cblowing ratioxe2x80x9d. The cooling air forms a protective xe2x80x9cfilmxe2x80x9d between the metal surface and the hot gas flow, preventing melting or other degradation of the component.
Film cooling performance may be characterized in several ways. One relevant indication of performance is known as the adiabatic wall film cooling effectiveness, sometimes referred to herein as the xe2x80x9ccooling effectivenessxe2x80x9d. This particular parameter is equivalent to the concentration of film cooling fluid at the surface being cooled. In general, the greater the cooling effectiveness, the more efficiently can the surface be cooled.
Under certain conditions, the cooling stream moving through a passageway in a substrate and out to the hot surface tends to separate from the hot surface quickly, rather than moving along the surface and being in close contact therewith. This separation can seriously diminish the cooling effectiveness and lead to temperature-related damage to the part. This is often the case with aircraft engine components, such as combustor liners, where the blowing ratio is usually greater than about 1, and often in the range of about 2 to about 6.
A reduction in the blowing ratio, e.g., to a preferable value less than 1, might help to prevent the cooling stream from separating from the surface. However, for most combustors, the velocity of the cooling stream is determined in large part by the pressure drop across the combustor liner, and the turbine designer is usually not able to significantly alter the blowing ratio without changing other critical parameters in the engine design. Moreover, the use of greater amounts of cooling air to try to maintain a certain cooling capacity diverts air away from the combustion zone. This can lead to other problems, such as greater air pollution resulting from non-ideal combustion, and less efficient engine operation.
One can readily understand that new methods for increasing the cooling effectiveness provided by a discrete hole film cooling system would be welcome in the art. The methods should be especially applicable to parts exposed to very high operating temperatures, such as metal-based turbine engine parts. Moreover, the discovered techniques should not interfere with other functions, e.g., the efficient operation of a turbine engine, or the strength and integrity of turbine engine parts. The methods should also be compatible with other protective systems which may be used simultaneously, such as thermal barrier coating systems. Finally, the implementation of these methods should preferably not involve bstantial cost increase in the manufacture or use of the relevant component, or of a system in which the component operates.
The needs discussed above have been met by the discoveries outlined herein. One embodiment of this invention is directed to a method for improving the cooling effectiveness of a gaseous coolant stream which flows through at least one passage hole in a substrate to an exit site on a high-temperature surface of the substrate. The method comprises disrupting the coolant stream at the exit site, so that the coolant stream contacts a greater area of the high-temperature surface. In many embodiments, this invention allows an increase in the concentration of gaseous coolant at the high-temperature surface by a multiplicative factor of at least about 1.1, as compared to a conventional coolant stream, and often, by a multiplicative factor of at least about 1.5.
In preferred embodiments, the exit site is a crater. Moreover, the exit site may be formed within a coating applied over the substrate, such as a thermal barrier coating.
Another embodiment of the present invention is directed to an article which comprises a substrate and at least one passage hole for a coolant stream extending through the substrate from a first surface to an exit site at a second surface which is selectively exposed to high temperature. The passage hole has a substantially uniform cross-sectional area within the substrate, but has a different cross-sectional area at the exit site, suitable for disrupting the flow of the coolant stream. The exit site may be in the shape of a crater, as discussed below. Moreover, the exit site may be located in a coating situated on top of the substrate.
One example of an article based on embodiments of the present invention is a metal-based component of a turbine engine, such as a combustor. As demonstrated below, this invention significantly increases the cooling effectiveness of a coolant stream typically employed to protect turbine engine components from excessive exposure to high temperature.